40mN/3000s氙离子电推进系统工作性能在轨测试与分析

为获得电推进系统准确的工作性能,验证其在轨工作稳定性,国内首台空间用氙离子电推进系统于2012年11月~2014年2月在我国新技术试验卫星——实践九号A卫星上开展了首次在轨飞行试验。本文简要介绍了40mN/3 000 s氙离子电推进系统的设计方案和主要性能指标,重点阐述了所开展的系统性能在轨测试内容、测试方法及获得的测试结果,并对测试结果进行了分析,最后给出了在轨测试结论。     

电推进系统空间试验技术研究

电推进系统的相容性、空间推力/比冲等是空间应用关注的重要性能指标。根据电推进系统未来空间试验技术发展趋势,调研了国内外离子、霍尔电推进系统的推力、电磁兼容性、对卫星的污染等空间试验情况,结合我国电推进系统首次开展空间试验现状、电推进系统的布局,以及星上配备的卫星污染与电位监测器,对空间环境条件下卫星的污染、电推进自身及卫星设备的电磁兼容性、空间推力标定方法、推进剂剩余量分析方法等进行了研究。通过电推进系统在轨连续试验、电推进羽流影响等分析,得到电推进对卫星周围污染情况、电推进与卫星平台的电磁兼容性等在轨性能参数,可为全面评价电推进系统技术、科学制定电推进空间试验计划及电推进空间应用提供依据。     

霍尔电推进流量调节模块研制及在轨验证

流量调节模块是霍尔电推进系统的关键模块,用于精确控制霍尔推力器的推进剂流量.上海空间推进研究所在多年研究工作的基础上,针对国内首次霍尔电推进空间飞行任务,选用金属多孔材料作为节流器件,成功研制了小型化微流量控制器及具有主备份支路的流量调节模块.本文介绍了流量调节模块设计思路、研制过程中的重要试验和在轨飞行验证情况,并针对后续电推进发展需求,提出流量调节模块的后续发展思路.     

嫦娥四号探测器推进系统设计特点与验证

嫦娥四号探测器推进系统在充分继承嫦娥三号探测器的技术方案基础上,在发动机精确控制和高精度变推力、轨道控制管路超压管理、月面高可靠推进剂钝化、气瓶在轨提高裕度和发动机量化极性测试方面进行了优化设计,并通过了地面试验和在轨飞行试验的充分验证。结果表明:嫦娥四号探测器推进系统完成了在轨飞行全部姿态控制和轨道控制任务,取得的发动机推力精确控制和高精度变推力等技术成果可应用于其他探测器。     

硬X射线调制望远镜卫星飞行程序设计及在轨验证

硬X射线调制望远镜(HXMT)卫星飞行程序设计具有在轨使用模式复杂、载荷需求约束多、在轨工作状态多且多系统多组合耦合、在轨自主管理功能应用与飞行事件操作流程耦合等特点。卫星在轨测试内容多、约束多、耦合强,因此在飞行程序设计中重点开展了入轨段程序时序评估和改进,观测模式、自主功能的应用设计,在分析设计需求约束并提取能够有效覆盖验证全部正常功能和工作状态的系统组合形式基础上,进行飞行事件编排优化。HXMT卫星在轨验证结果表明:卫星在轨运行良好,各飞行事件有序开展,飞行程序设计合理有效。     

层次分析法在空间推进系统综合评价中的应用

在建立航天器推进系统综合评定体系的基础上，动用层次分析法对已经应用飞行的推进系统和有未来应用潜力的推进系统进行了评估。评估结果表明，从能量特性，使用性能、系统可靠性、经济性、系统特性五个准则上看，过氧化氢／煤油无毒推进系统是未来航天器在轨推进的最优选择。     

离子电推进非预期电击穿的工程防护技术研究

离子电推进的非预期电击穿现象具有普遍性和复杂性,结合实例对非预期电击穿的现象和分类、击穿事件的一般处理方法、工程隐患等基本问题以及实践二十卫星上LIPS-300离子电推进在轨工作初期的击穿情况进行了简要介绍,对基于推力器及其工作环境、电推进系统及其航天工程的非预期电击穿控制与防护技术进行了全面系统的讨论,通过对推力器栅极材料、电场强度、束流密度和推力器环境中低气压、等离子体、多余物、溅射沉积物等因素的控制,以及在电推进电源处理单元健壮性、系统监测与响应、沉积能量控制等方面的改进,再加上合理的在轨处理程序,包括入轨预处理、连续击穿循环控制程序和栅间短路清除技术等措施,可有效降低击穿频次及其对电推进系统造成的影响。     

电推进航天器的特殊环境及其影响

文章介绍了应用离子和霍耳电推进系统在航天器周围产生的等离子体和电磁场等特殊环境,讨论了这些特殊环境对航天器各分系统或部件产生的溅射腐蚀、沉积污染、充放电、等离子体干扰、碰撞动力学扰动等影响效应,探讨了研究电推进与航天器相互作用效应的地面试验技术、空间飞行试验技术和模型分析技术,介绍了离子电推进系统与航天器相容性分析评价的技术要点。     

支持航天员操作的电连接器设计与验证

针对空间站建造期间在轨维修、航天员出舱等关键功能和任务技术难度大、空间环境复杂,常规电连接器在轨操作性、安全性及环境适应性不能满足系统功能需求的问题,文章提出了一种覆盖全系统研制阶段的航天员操作的电连接器设计与验证方法,以为复杂航天器关键功能器件的系统设计与应用提供参考。经空间站在轨验证,结果表明：运用该设计技术及验证方法的电连接器能够高效、可靠、安全地支持航天员开展在轨维修活动,可为载人航天器系统维修性设计提供技术支撑。     

科技成果

<正>X-37B将执行第4次飞行任务据腾讯网2015年5月11日报道,美国空军的X-37B航天飞机将进行第4次飞行,用于试验最新的在轨机动与任务延伸技术。此次任务涉及到美国空军研究实验室和美国航天与导弹系统中心联合开发的一种新型推进系统,X-37B将利用这套系统在轨机动,试验其基本性能,空军官员将评估X-37B的性能,并进行一些改进。经过前3次的飞行试验,X-37B已经累计在轨1366天,美国空军已经掌握了其飞行性能,有助于之后战斗力的形成,X-37B由波     

电推进羽流与航天器相互作用的研究现状与建议

电推进与传统的化学推进相比可以节约大量推进剂质量,被广泛用作地球同步轨道卫星南北位置保持和深空探测等任务的主推进系统。电推进器工作时产生的羽流与传统的化学推进器羽流有显著区别,电推进羽流对航天器的影响是进行卫星电推进器系统设计时需要重点关注的问题。文章讨论了电推进羽流对航天器的主要影响,介绍了国外在地面模拟试验、空间飞行验证和软件仿真技术等方面的研究现状,同时对国内开展羽流与航天器相互作用研究提出了建议。     

水基推进系统综述

水基推进系统的发展为航天器推进、供电、能量储存、系统构建提供了新的选择。这一系统的核心是一体化可再生燃料电池,它利用太阳能电池帆板提供的电能电解水．产生的氢氧气体既可以用于燃烧推进,又可以通过燃料电池反应重新复合为水进行供电。介绍了水基推进系统的基本组成与工作原理,通过质量估算与性能分析,评估了水基推进系统的工作能力、适用范围及空间应用优势。     

On a possibility of delivery of samples of primordial matter from asteroids of the main belt by a spacecraft with electric propulsion

The possibilities of an expedition to asteroids of the Main belt with the goal of obtaining samples of their matter using a spacecraft with a small-thrust electric propulsion, which is under development as a part of the Fobos-Grunt project, are considered. Obtaining matter from different regions of the Solar system is necessary in order to understand its origin and the structure of the Earth. Approximate estimates are made for the flights of spacecraft with electric propulsion and different power of solar batteries, and with perturbation maneuver near Mars for expeditions to the Main asteroid belt, including its middle part.     

基于电火箭推进的地-月轨道转移自主导航研究

以电推进力作为奔月航天器在进入月球影响球之后的螺旋式减速转移过程的主推进力和姿态控制力,建立限制性三体力学模型.采用星光导航方式,结合UKF非线性滤波方法进行自主导航研究.推力的引入方式能有效抑制由于航天器姿态与导航误差耦合而导致的推力矢量方向误差.通过大量的计算机仿真,详细研究和分析了电推进条件下自主导航的发散问题,有针对性地给出了抑止发散的方法,为工程实现提供了有利的参考依据.     

Use of the Lyapunov Functions for Calculating the Locally Optimal Control of a Thrust Vector during Low-Thrust Interorbital Transfer

Cosmic Research - This paper presents a method for locally optimal control of the thrust vector of the electric propulsion system (EPS) for a spacecraft that performs a multiturn interorbital...     

Design of optimal earth pole-sitter transfers using low-thrust propulsion

Recent studies have shown the feasibility of an Earth pole-sitter mission using low-thrust propulsion. This mission concept involves a spacecraft following the Earth's polar axis to have a continuous, hemispherical view of one of the Earth's poles. Such a view will enhance future Earth observation and telecommunications for high latitude and polar regions. To assess the accessibility of the pole-sitter orbit, this paper investigates optimum Earth pole-sitter transfers employing low-thrust propulsion. A launch from low Earth orbit (LEO) by a Soyuz Fregat upper stage is assumed after which solar electric propulsion is used to transfer the spacecraft to the pole-sitter orbit. The objective is to minimize the mass in LEO for a given spacecraft mass to be inserted into the pole-sitter orbit. The results are compared with a ballistic transfer that exploits manifold-like trajectories that wind onto the pole-sitter orbit. It is shown that, with respect to the ballistic case, low-thrust propulsion can achieve significant mass savings in excess of 200 kg for a pole-sitter spacecraft of 1000 kg upon insertion. To finally obtain a full low-thrust transfer from LEO up to the pole-sitter orbit, the Fregat launch is replaced by a low-thrust, minimum time spiral, which provides further mass savings, but at the cost of an increased time of flight.     

Optimization of the trajectory of the spacecraft insertion into the system of heliocentric orbits

The possibility of the spacecraft insertion into the system of operational heliocentric orbits has been analyzed. It has been proposed to use a system of several operational heliocentric orbits. On each orbit, the spacecraft makes one or more revolutions around the Sun. These orbits are characterized by a relatively small perihelion radius and relatively high inclination, which allows one to investigate the polar regions of the Sun. The transition of the spacecraft from one orbit to another has been performed using an unpowered gravity assist maneuver near Venus and does not require the cruise propulsion operation. Each maneuver transfers the spacecraft into the sequence of operational heliocentric orbits. We have analyzed several systems of operational heliocentric orbits into which the spacecraft can be inserted by means of the considered transportation system with electric propulsion (EP). The mass of the spacecraft delivered to these systems of operational orbits has been estimated.     

Project of the space test&fitting-out dock for solar-powered electric propulsion on-board international space station

A concept of a universal fitting-out module applicable for space flight test of a solar-powered electric propulsion (EP) as well as for the assembly and deployment of small spacecraft (S/C) from the international space station (ISS) is proposed.     

Effect of thin condensate films of a metal working fluid of an electric propulsion engine on the integral optical coefficients of a spacecraft’s thermal control coating

Materials on experimental studies to determine the effect of thin condensate films of cesium (used as a model working medium for electric propulsion engines and some spacecraft power sources) on integral optical coefficients of spacecraft thermal control coatings are presented. A technique modified by the author and employing the regular thermal regime of a thin metal plate is used. Measurement results demonstrate that films with thicknesses of 100–1000 Å can seriously degrade the integral optical coefficients of thermal control coatings and thus disturb the heat balance of some spacecraft systems.